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Publication numberUS7699269 B2
Publication typeGrant
Application numberUS 12/193,213
Publication dateApr 20, 2010
Filing dateAug 18, 2008
Priority dateMar 11, 2004
Fee statusPaid
Also published asUS7424989, US20050224659, US20090200430
Publication number12193213, 193213, US 7699269 B2, US 7699269B2, US-B2-7699269, US7699269 B2, US7699269B2
InventorsDale M. Pitt
Original AssigneeThe Boeing Company
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Intelligent multifunctional actuation system for vibration and buffet suppression
US 7699269 B2
Abstract
Circuits and methods for use on a mobile platform including a flight control system, a structure, an aerodynamic surface, and an actuator operatively coupled to the surface to control the surface. The circuit includes a first and a second input, a summing element, and an output. The first input accepts commands from the flight control system while the second input accepts a vibration signal from the structure. The summing element communicates with the inputs and sums the signal and the command. In turn, the summing element controls the actuator with the summed signal and command.
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Claims(17)
1. A mobile platform comprising:
a flight control system generating commands;
a structure associated with the mobile platform;
an aerodynamic surface;
an actuator operably coupled to the surface to control the surface;
a vibration sensor coupled to the structure and generating a vibration signal representative of a vibration of the structure;
a position sensor that generates a position signal representative of the position of the actuator, the flight control system being responsive to the position signal; and
a circuit in communication with the actuator, the position sensor and the vibration sensor, with the circuit including;
a first input communicating with the flight control system to accept the commands;
a second input communicating with the vibration sensor to accept the vibration signal,
a third input to receive the position signal indicating the position of the actuator; and
a summing element that sums the vibration signal and the command, the circuit controlling the actuator with the summed signal and command.
2. The mobile platform according to claim 1, further comprising a filter in communication with the third input associated with the position sensor, and the vibration sensor, the filter filtering the vibration signal from the position signal, the filter including an output and communicating the filtered position signal to the flight control system via the output.
3. The mobile platform according to claim 1, further comprising a first power cable connected to the actuator and supplying power to the actuator, the actuator including a second power cable communicating with the circuit and the first power cable.
4. The mobile platform according to claim 1, wherein the mobile platform is an aircraft.
5. The mobile platform according to claim 4, wherein the structure is a portion of a wing.
6. The mobile platform according to claim 4, wherein the structure is a housing of the actuator.
7. The mobile platform according to claim 4, wherein the surface is an aileron.
8. The mobile platform according to claim 7, wherein the actuator is at an outboard location of the aileron.
9. A mobile platform comprising:
a control system generating commands;
a structure;
an aerodynamic surface;
an actuator operably coupled to the aerodynamic surface to control the aerodynamic surface;
a vibration sensor coupled to the structure and generating a vibration signal representative of a vibration of the structure; and
a circuit coupled to the actuator adapted to communicate with the control system and the vibration sensor and to use the commands and the vibration signal to control the actuator;
wherein the circuit includes:
a first input communicating with the control system to accept the commands;
a second input communicating with the vibration sensor to accept the vibration signal;
a summing element communicating with the inputs to sum the signal and the command;
a third input for accepting a position signal representative of a position of the actuator; and
a filter in communication with the second and third inputs adapted to filter the vibration signal from the third signal.
10. The mobile platform of claim 9, wherein the mobile platform comprises an airborne mobile platform.
11. The mobile platform of claim 9, wherein the mobile platform comprises an aircraft.
12. The mobile platform of claim 9, further comprising a first power cable connected to the actuator and supplying power to the actuator.
13. The mobile platform of claim 12, wherein the actuator includes a second power cable communicating with the circuit and the first power cable.
14. The mobile platform of claim 9, wherein the structure comprises a wing.
15. An airborne mobile platform comprising:
a flight control system generating commands;
a structure that forms a portion of the airborne mobile platform;
an aerodynamic surface;
an actuator operably coupled to the aerodynamic surface to control the aerodynamic surface;
a vibration sensor coupled to the structure and generating a vibration signal representative of a vibration of the structure; and
a circuit coupled to the actuator adapted to communicate with the control system and the vibration sensor and to use the commands and the vibration signal to control the actuator; and
the circuit including:
a position sensor in communication with the vibration sensor that senses a position of the actuator;
a filter in communication with the position sensor and the vibration sensor, and adapted to filter the vibration signal from the position signal, and to output a filtered position signal to control said flight control system.
16. The airborne mobile platform of claim 15, wherein the structure is a portion of a wing.
17. The airborne mobile platform of claim 15, wherein the surface is an aileron.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. patent application Ser. No. 10/798,687 filed on Mar. 11, 2004. The entire disclosure of the above application is incorporated herein by reference.

FIELD

This disclosure relates generally to flight control systems and, more particularly, to flight control systems for suppressing aerodynamically induced vibrations.

BACKGROUND

High performance fighter aircraft such as the F-18 and F-22 (available from the Boeing Company of Chicago, Ill.) often experience high frequency aerodynamically induced vibrations of their wings, stabilators, and vertical tails. These vibrations are cause by buffeting aerodynamic forces and are transmitted into, and through, the aircraft structure. If uncompensated for, the associated stresses may lead to premature cracking of the structure. The resulting repairs of these cracks are both expensive and time consuming. In the alternative, aerodynamic modifications to reduce the causative turbulence impose performance constraints on the aircraft while structural modifications to reduce the resulting fatigue stress incur weight and cost penalties.

Additionally, the buffeting of the vehicle transmits noise into, and causes noise within, the aircraft structure. In turn, the structure transmits the noise to the cockpit wherein noise control measures, with additional penalties must be undertaken. Nor are these problems isolated to high performance aircraft. Automobiles, missiles, and launch and reentry vehicles (for example) also receive buffeting from aerodynamic forces.

Various attempts have been made to use the existing flight control actuators to compensate for these aero-vibrations. However, the flight control system typically commands the actuator at rates of about 30 cycles per second or less. Since the vibrations occur at frequencies significantly higher than the commands, such attempts have failed.

Moreover, attempts to modify the flight control system to react quickly enough to respond to the vibration are impractical for a variety of reasons. For instance, such modifications require an order of magnitude increase in the flight computer speed. Thus, these solutions necessitate an upgrade of the computer. Additionally, modifying the flight control laws to accommodate the additional functionality necessitate the recertification of the flight control system. These recertifications are expensive, time consuming, and (as such) highly undesirable.

Furthermore, as composite materials replace aluminum, and other conventional, structural members (e.g. on the Boeing 7E7 aircraft) vibration control may assume an increasing importance in the design, operation, and maintenance of aircraft. Thus a need exists to reduce or eliminate aerodynamically induced vibrations.

SUMMARY

It is in view of the above problems that the present disclosure was developed. The disclosure provides apparatus and methods for reducing vibrations of mobile platform structures.

In one embodiment, the disclosure provides a self-contained actuation device that reduces the aerodynamic buffet loads. Accordingly, construction and operation of mobile platforms (e.g. aircraft) in accordance with the principles of the present disclosure results in more cost efficient platforms that possess better performance and longer service lives.

Another embodiment employs the existing flight control actuators to reduce the aero-buffeting without requiring modification of the flight control system. The present embodiment also includes a vibration sensor placed on an aircraft structure (e.g. a wing) to sense the aero-vibration. The sensor is connected to a controller that is coupled to the actuator. The controller inverts the vibration signal and adds it to the actuator command from the flight control system. Then the controller sends the combined signal (the command with the inverted vibration signal superimposed thereon) to the actuator, thereby driving the actuator out of phase with the vibration. By thus canceling the vibration the current embodiment reduces cyclical loads and improves the fatigue life of the structure.

In yet another embodiment, a method of reducing aerodynamically induced vibrations is provided. The method includes sensing the vibration and inverting a signal representative of the same. The inverted vibration signal is superimposed on a flight control system command for an actuator to drive the actuator out of phase with the vibration.

Further features and advantages of the present disclosure, as well as the structure and operation of various embodiments of the present disclosure, are described in detail below with reference to the accompanying drawings.

DRAWINGS

The accompanying drawings, which are incorporated in and form a part of the specification, illustrate the embodiments of the present disclosure and together with the description, serve to explain the principles of the disclosure. In the drawings:

FIG. 1 illustrates an aircraft constructed in accordance with the principles of the present disclosure;

FIG. 2 is a detailed view of a wing of the aircraft of FIG. 1;

FIG. 3 shows an actuator in accordance with the principles of a embodiment of the present disclosure;

FIG. 4 is a block diagram of a control system in accordance with the principles of another embodiment of the present invention;

FIG. 5 is a schematic diagram of a circuit in accordance with a embodiment of the present disclosure; and

FIG. 6 is a flowchart of a method in accordance with a preferred form of the present disclosure.

DETAILED DESCRIPTION

Referring to the accompanying drawings in which like reference numbers indicate like elements, FIG. 1 illustrates a pair of aircraft 10 constructed in accordance with an embodiment of the present disclosure. The aircraft 10 includes a number of structures and control surfaces that are well known in the art and will herein be represented by an exemplary wing 12 and an aileron 14. Additionally, a fairing 18 that covers an aileron actuator may be seen on the underside of the wing 12.

During flight, the flight computer and the pilot continually sense the flight related conditions (pitch, climb rate, speed, and the like). Depending on the conditions, either the computer, the pilot, or both issue commands (e.g. electro-magnetic signals) to reposition the ailerons 14 and other control surfaces. These commands cause the actuator 20 to either extend or retract to move the aileron 14. As noted previously, the wing 12 rushing through the air causes turbulence in the air that causes the wing 12 to vibrate. Additionally, the movement of the control surface aileron 14 tends to change the airflow, thus introducing another source of turbulence, buffeting, and vibration.

Referring now to FIG. 3, an actuator 20 for the aileron 14 is shown. The actuator 20 includes an attachment fitting 22 for coupling the actuator 20 to the wing 12 and a ram 24 that moves in response to flight control system commands. The ram 24 is operatively coupled to the aileron 14 to change the aileron position. Additionally, FIG. 3 shows a circuit 26 being mechanically coupled to the actuator 20 as well as a vibration sensor 28 rigidly coupled to the actuator 20. While FIG. 3 shows the vibration sensor 28 separate from the circuit 26, the sensor 28 may be included in the circuit 26. It should also be noted that the actuator 20 may also include an internal position sensor such as an LVDT (linear variable differential transformer) to detect the position of the ram 24.

Additionally, the actuator typically includes a conventional power cable 30 for receiving power from the aircraft 10 power subsystem. A second power cable 32 may branch from the first power cable 30 to supply the circuit 26 power. Importantly, branching the second power cable 32 from cable 30 saves cable weight by obviating the need for a dedicated cable run from the aircraft 10 power subsystem to power the circuit 26. Another cable 30′ carries the flight control system command to the circuit 26, while in previous systems the cable 30′ was connected directly to the actuator instead of the circuit.

Moreover, coupling the vibration sensor 28 to the actuator 20 instead of the wing 12 (FIG. 1) eliminates the need for, and associated weight of, a vibration signal cable from the wing to the circuit 26. Though, the vibration sensor 28 may be located on the wing 12 without departing from the spirit or scope of the disclosure. Additionally, the elimination of the wing-to-circuit vibration signal cable simplifies the interfaces with the wing since a cable 31 on the actuator 20 suffices to carry the vibration signal to the circuit 26. In turn, the simplification reduces airframe integration costs. Similarly, integrating the vibration sensor 28 with the circuit 26, rather than mounting the two on the actuator 20 separately, simplifies construction and installation of the actuator and reduces costs further still.

Furthermore, coupling the circuit 26 to the actuator 20 eliminates the need for cables between the circuit and the actuator to carry the command signal from the circuit and to carry the position signal from the actuator. Likewise, because the flight control system already includes a cable 30′ to the actuator (to carry the actuator command) the present disclosure requires no modifications to the flight control system. Accordingly, the present disclosure further reduces integration costs and delays.

With reference now to FIG. 4, the block diagram shown therein represents a system in accordance with an exemplary embodiment of the present disclosure. The system 100 includes a summing element 102, an actuator, 104, a control surface 106, a structure 108 (e.g. a wing), a vibration sensor 110, an inverter 112, a filter 114, and an actuator position sensor 116. Upon arrival from the flight control system a command passes through the summing element 102 (assuming for the moment there is no current vibration). The command causes the actuator 104 to move and reposition the control surface 106. Because of aerodynamic turbulence over the control surface 106 and wing 108, the wing begins to vibrate. In turn, the vibration sensor 110 generates a signal representative of the vibration that the inverter 112 inverts.

To complete the feedback loop, the summing element 102 superimposes the inverted vibration signal on the command. The resulting signal (the command with the inverted vibration signal superimposed on it) is fed to the actuator 104 to drive the control surface out of phase with the vibration. Accordingly, the control surface 106 causes a disturbance that cancels the vibration of the wing 108. Those skilled in the art will recognize that the commands typically operate up to approximately 30 Hz while the vibration occurs at substantially higher frequencies. Similarly, the commands are typically signals having amplitudes well in excess of the amplitude of the vibration signal (or can be so tailored with appropriate choice of system gains). Thus the inverted vibration signal appears as a ripple superimposed upon the command signal after the two are added.

Since many flight control systems are designed to sense the actual position of the actuator 104, a feedback signal may also be provided by the present disclosure. In particular the vibration sensor 110 and the position sensor 116 (of the actuator 104) may communicate with the filter 114 that subtracts the vibration signal from the position signal as sensed by the position sensor. Accordingly, the position signal fed back to the flight control system does not reflect the slight difference between the commanded position and the actual position as influenced by the vibration canceling circuit 126. Thus, by filtering the vibration signal from the actuator position signal, the present disclosure ensures that the flight control system operates properly (e.g. does not raise a false alarm to indicate that the actuator is out of position).

Note should also be made that FIG. 4 depicts a circuit 126. The circuit 126 includes the summing element 102, the inverter 112, and perhaps the filter 114 connected as shown. Hence, the circuit 126 combines the command and the inverted vibration signal to cancel the vibration. The circuit 126 may also provide the filtered position signal as shown in FIG. 4. In one embodiment, the control surface 106 is the aileron 14 (FIG. 1) and the vibration sensor is placed as close to the wing tip 12 as is practical. In this manner, the flexibility of the wing 12 and the position of the sensor 110, well out on the wing ensure that a robust vibration signal is generated.

Now with reference to FIG. 5, a simplified schematic of a circuit 226 in accordance with an embodiment of the present disclosure is illustrated. The circuit 226 includes a summing amplifier 202, an inverting amplifier 204, a difference amplifier 206, and an internal power supply 208. Interfaces with the circuit 226 include a flight control system command input 210, a vibration signal input (here, the vibration sensor 212 is shown as being internal so that the interface is not necessary), a command output 214, an actuator position signal input 216, and a position signal output 218. Internally, an inverted vibration signal 220 is also shown.

Additionally, a flight control actuator 228 is shown external to the circuit 226. Schematically FIG. 5 illustrates the actuator 228 as a solenoid to indicate that the actuator contains components that require electric power. The required electric power is typically supplied by the aircraft power subsystem 224 via a cable 230. A second power cable 232 is shown branching from the cable 230 and providing power to the power input 222 for the circuit 226. Thus, the circuit 226 and actuator share a common power supply. Those skilled in the art will recognize that the internal power supply 208 is shown in simplified form for clarity. Details such as neutral returns and interconnections to the amplifiers 202 to 206 have been likewise omitted.

Additionally, a housing 234 of the circuit 226 is rigidly coupled to the actuator so that the vibration transducer 212 accurately senses the vibration of the actuator 228. In turn, the actuator 228 is rigidly mounted to the structure for which vibration reduction is desired (e.g. by the attachment means 22 shown by FIG. 3). Because of the rigid coupling between the structure and the actuator 228, and between the actuator and the sensor 212, the sensor 212 provides a reliable indication of the vibration of the structure.

Thus, in operation, the summing amplifier 202 sums the flight control command 210 and the vibration signal 220 as inverted by the inverting amplifier 204. The summing amplifier 202 outputs the command 214, with the inverted vibration signal superimposed thereon, to the actuator. As noted previously, the actuator 228 causes an aerodynamic disturbance that cancels the vibration of the structure. Subsequently, if another transient disturbance causes the vibration to return, the vibration cancellation loop (as just described) acts to cancel the new vibration. In the alternative, a non-inverted vibration signal could be subtracted from the command to the same general effect of canceling the vibration.

The circuit 226 also includes a position feedback subsystem that includes the difference amplifier 206. The amplifier 206 accepts the position signal 216 from the actuator position sensor and subtracts the position signal 216 from the vibration signal supplied by the sensor 212. Accordingly, the amplifier 206 filters the vibration signal from the position signal 216 and communicates the resulting signal 218 to the flight control system.

With reference now to FIG. 6, a flowchart illustrating a method in accordance with the principles of the present disclosure is shown. The method includes mounting a vibration canceling circuit to either the wing 12 or the aileron actuator as shown at 302 and 304 respectively. In operations 306 and 308 a vibration sensor is coupled to either the wing 12 or the actuator. Jumpers are then installed between the circuit and the vibration sensor, the circuit and a position sensor of the actuator, and the circuit and the actuator power supply (preferably within the envelope of the actuator). See operation 310. Moreover, the flight control system cable that carries the actuator command is moved from the actuator to the circuit in operation 311.

With the power on, commands for the actuator may then be received while the vibration is sensed as at 312 and 314. The two signals are added, subtracted, or superimposed to obtain a command with an inverted vibration signal superimposed upon it at operation 315. The combined signal, from operation 315, is used to drive the actuator out of phase with the vibration. See operation 316. Meanwhile, the position of the actuator may be sensed in 318. Accordingly, the vibration signal may be filtered from the position feedback signal and the result forwarded the flight control system. See for instance operation 320. In this manner, the method may repeat for as long as vibration canceling is desired as indicated at decision 322.

In view of the foregoing, it will be seen that the several advantages of the disclosure are achieved and attained. Notably, the embodiments described herein require no added cabling, or cabling modifications outside of the envelope of the actuator. Moreover, the present disclosure requires no modification to the flight control system, or even the flight computer. Thus, the present disclosure provides vibration elimination with a weight and cost savings over previous attempts to reduce aerodynamically induced vibrations.

The embodiments were chosen and described in order to best explain the principles of the disclosure and its practical application to thereby enable others skilled in the art to best utilize the disclosure in various embodiments and with various modifications as are suited to the particular use contemplated. For example, many industrial machines include structures that are moved by an actuator commanded by a control system. If the machine includes a tool (e.g. a drill) on the movable structure, the tool may induce undesired vibrations in the structure. Thus, the apparatus and methods discussed herein may be adapted to sense the tool induced vibration of the structure and cancel the same without requiring modification of the machine control system. Accordingly, machines with increased precision and accuracy are also provided by the present disclosure.

As various modifications could be made in the constructions and methods herein described and illustrated without departing from the scope of the disclosure, it is intended that all matter contained in the foregoing description or shown in the accompanying drawings shall be interpreted as illustrative rather than limiting. Thus, the breadth and scope of the present disclosure should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims appended hereto and their equivalents.

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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US8090485 *Nov 27, 2007Jan 3, 2012Embraer S.A.Low-frequency flight control system oscillatory faults prevention via horizontal and vertical tail load monitors
US20090138147 *Nov 27, 2007May 28, 2009Erick Vile GrinitsLow-frequency flight control system oscillatory faults prevention via horizontal and vertical tail load monitors
Classifications
U.S. Classification244/174, 244/99.13
International ClassificationB64C13/00, B64D1/12, B64C13/16
Cooperative ClassificationB64C13/16, Y02T50/44
European ClassificationB64C13/16
Legal Events
DateCodeEventDescription
Aug 19, 2008ASAssignment
Owner name: THE BOEING COMPANY, ILLINOIS
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:PITT, DALE M.;REEL/FRAME:021410/0510
Effective date: 20040305
Owner name: THE BOEING COMPANY,ILLINOIS
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:PITT, DALE M.;REEL/FRAME:021410/0510
Effective date: 20040305
Oct 21, 2013FPAYFee payment
Year of fee payment: 4
Oct 20, 2017MAFP
Free format text: PAYMENT OF MAINTENANCE FEE, 8TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1552)
Year of fee payment: 8